1. Technical field
The invention relates to a process for fabrication of parts with large dimensions comprising a skin and stiffeners, and a composite material with a thermoplastic matrix.
The process according to the invention may be used in many industrial sectors, if it is required to benefit from the inherent advantages of composite materials for making parts with large dimensions comprising a skin and add-on stiffeners. Thus in the aeronautical industry, the process according to the invention may be used in particular to make aircraft fuselage segments, aircraft jet engine shrouds, etc.
2. State of the Art
When it is required to make a thin envelope with good mechanical behavior but without excessively increasing its weight, that will be called a xe2x80x9cskinxe2x80x9d throughout the rest of the text, a frequent means of achieving this end is to use stiffeners such as rails, frames, local stiffeners, etc., added onto the skin.
In the past, this type of structure was always fully metallic, the skin being in the form of a sheet metal plate and the stiffeners being in the form of sheet metal plates or sections added onto the skin by fasteners such as rivets. Many metallic structures are still made in this way, particularly for parts with large dimensions. In particular in the aeronautical industry, aircraft fuselage segments and jet engine shrouds are always made in this way, particularly as described in documents US-A-5 560 102 and US-A5 586 381
In the last few years, an increasingly large number of metallic parts are being replaced by composite material parts formed of long fibers such as carbon fibers embedded in a resin matrix. This change is explained by the advantages specific to composite materials. In particular, these advantages include a weight saving of about 25% compared with comparable metallic parts, while the mechanical properties are similar to the properties of metallic parts and can be modified on request. Composite material parts also have good resistance to fatigue, no corrosion and excellent specific properties. The weight saving and the excellent mechanical properties of composite materials with long fibers embedded in a resin matrix explain the outstanding penetration of these materials in the aeronautical industry.
More precisely, the vast majority of composite material parts now used in the aeronautical industry are made from a thermosetting resin. Initially, this was the only type of resin that was capable of providing the required mechanical properties.
However, the fabrication of parts made from composite materials with a thermosetting resin has a significant disadvantage. Due to the thermosetting nature of the resin used, the fabrication of each part necessarily terminates with a relatively long polymerization operation, usually carried out in an autoclave.
This final operation is not particularly penalizing in the case of small parts. Small and relatively inexpensive autoclaves can be used and several parts can be polymerized in each autoclave at the same time.
On the other hand, when the size of parts is larger, only one part can be polymerized at any one time in the same autoclave, and very large and very expensive autoclaves must be used. The duration of the operation and the cost of the autoclave then quickly make the process unsuitable for industrial applications. This is why, although documents US-A-5 170 967 and US-A-5 223 067 envisage the use of a composite material with a thermosetting matrix for the fabrication of an aircraft fuselage segment, it is difficult to justify industrial production of large parts using this technology.
Since the relatively recent appearance of thermoplastic resins such as PEEK (Polyetheretherketone) resin that can be used to make composite materials with long fibers and with a thermoplastic matrix and that have mechanical properties equivalent to the mechanical properties of the most recent composite materials with a thermosetting matrix, there is a trend towards replacing existing metallic parts by parts made using a composite material with a thermoplastic matrix.
Apart from the advantages of composite materials with an organic matrix mentioned above, these composite materials with a thermoplastic matrix have good resistance to impact and fire and low moisture absorption. Finally, partly finished products can be kept at ambient temperature and have a practically unlimited life due to the fact that the resin that impregnates the threads is already polymerized.
As illustrated particularly in document US-A-5 362 347, it has already been suggested that the leading edge of an aircraft wing can be made using a composite material with a thermoplastic matrix. More precisely, this document describes that stiffeners and the skin can be made separately and then assembled by welding/diffusion.
Due to the fact that the stiffeners and the skin are made separately before being assembled, consolidation operations must be performed on the skin andxe2x80x94the stiffeners before they are assembled together. Remember that the main function of these consolidation operations is to create a bond between the various layers that form firstly the skin, and secondly each of the stiffeners, while eliminating pores. They consist of applying pressure on the element to be consolidated and heating it to a given temperature greater than the melting temperature of the resin. These consolidation operations are carried out in an autoclave. This makes this process unsuitable for the fabrication of parts with a skin with large dimensions, for the same reasons as mentioned above in the discussion for the fabrication of parts made of a composite material with a thermosetting matrix.
The purpose of this invention is precisely a process for making parts with large dimensions, such as aircraft fuselage segments, using a composite material with a thermoplastic matrix, in a particularly fast and inexpensive manner suitable for industrial fabrication at a relatively high rate, without any real limitation on the size due to the fact that neither the skin or the final structure obtained needs to be placed in an autoclave at the end of fabrication.
According to the invention, this result is achieved by means of a fabrication process for parts-with large dimensions, from a composite material comprising a skin and stiffeners, characterized by the fact that it comprises the following steps:
separate fabrication of stiffeners by lay-up, consolidation and shaping, starting from a strip of long fibers impregnated with thermoplastic resin;
placement of stiffeners on a tooling with a shape complementary to the shape of the part to be fabricated; and
fabrication of the skin and simultaneous assembly of the skin and the stiffeners, by lay-up and continuous consolidation of at least one strip of long fibers impregnated with thermoplastic resin, directly on the tooling on which the stiffeners are placed.
Due to the fact that the stiffeners and the skin are assembled at the time of fabrication of the skin, and due to the fact that this fabrication includes continuous consolidation of a strip of fibers forming the skin at the time of its lay-up, parts with arbitrary dimensions can be made without it being necessary to place the skin or the part obtained in an autoclave.
Furthermore, the fabrication time for parts is particularly short, since the skin and stiffeners are assembled at the same time as the skin is fabricated. Furthermore, stiffeners in a given part may be made while the skin of the previous part is being fabricated.
In one preferred embodiment of the invention, stiffeners are made by continuously and automatically depositing a strip of long fibers impregnated with polymerized thermoplastic resin in order to form a panel, similar to a piece of sheet metallic in a conventional metallic part, by cutting blanks out of this panel, and then by consolidation and shaping of these blanks.
In general, the blanks are consolidated before they are shaped.
The blanks may be consolidated either in an autoclave or in a heating press. Note that consolidation in an autoclave or in a heating press then concerns relatively small parts, such that an ordinary autoclave or an ordinary heating press with conventional dimensions can be used and several blanks can be consolidated in the press simultaneously.
Furthermore, the blanks are preferably shaped by thermoforming. Thermoforming may be preceded by a bending operation if the fibers need to be well oriented in the stiffener.
In some particular cases, and particularly when the stiffeners have a relatively limited curvature compared with a plane, the blanks may be consolidated and shaped simultaneously by thermoforming in a heating press.
The process according to the invention is advantageously applied to the fabrication of a hollow part. Stiffeners are then placed in recesses formed on a mandrel forming part of the tooling, the lay-up is then done and the strip is continuously consolidated on this mandrel by rotating it around its axis.
In the special case in which the process according to the invention is applied to the fabrication of aircraft fuselage segments, stiffeners comprising rails, frames and local stiffeners are fabricated.